Seal ring assembly for a gas turbine engine

ABSTRACT

A rotor assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a rotor that has a hub carrying one or more rotatable blades. The rotor is mechanically attached to a shaft, and an annular seal is carried by the shaft. The annular seal includes a substrate, a first layer disposed on the substrate, and a second layer disposed on the first layer and arranged to establish a sealing relationship with the rotor. The second layer includes a solid lubricant that has molybdenum trioxide (MoO3). A method of sealing for a gas turbine engine is also disclosed.

BACKGROUND

This disclosure relates to sealing components of a gas turbine engine.

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section, and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section.

The compressor section can include rotors that carry airfoils tocompress the air entering the compressor section. A shaft may be coupledto the rotors to rotate the airfoils.

SUMMARY

A rotor assembly for a gas turbine engine according to an example of thepresent disclosure includes a rotor that has a hub carrying one or morerotatable blades. The rotor is mechanically attached to a shaft, and anannular seal is carried by the shaft. The annular seal includes asubstrate, a first layer disposed on the substrate, and a second layerdisposed on the first layer and arranged to establish a sealingrelationship with the rotor. The second layer includes a solid lubricantthat has molybdenum trioxide (MoO3).

In a further embodiment of any of the foregoing embodiments, the firstlayer comprises copper.

In a further embodiment of any of the foregoing embodiments, the firstlayer has a composition, by weight percent, 60% to 95% copper, andaluminum.

In a further embodiment of any of the foregoing embodiments, the firstlayer has a composition, by weight percent, 8% to 12% aluminum, and atleast 85% copper.

In a further embodiment of any of the foregoing embodiments, thesubstrate comprises a nickel alloy.

In a further embodiment of any of the foregoing embodiments, the rotoris an integrally bladed rotor, and the annular seal extends about anouter diameter portion of the shaft.

In a further embodiment of any of the foregoing embodiments, the firstlayer has a composition, by weight percent, 85% to 95% copper, and atleast 5% aluminum.

In a further embodiment of any of the foregoing embodiments, the annularseal is a split ring including a seal body extending between a first endand a second, opposed end that engages with the first end.

In a further embodiment of any of the foregoing embodiments, theintegrally bladed rotor is a compressor rotor, and the substrate of theannular seal is seated in an annular groove extending inwardly from theouter diameter portion of the shaft, and the second layer abuts againstan inner diameter portion of the hub.

In a further embodiment of any of the foregoing embodiments, themolybdenum trioxide (MoO3) is formed from molybdenum disulfide (MoS2).

A gas turbine engine according to an example of the present disclosureincludes a fan section that has a fan, a compressor section that has acompressor, a turbine section that has a turbine coupled to thecompressor, and a shaft rotatable about an engine longitudinal axis. Thecompressor section includes a rotor assembly. The rotor assemblyincludes a rotor that has a hub carrying a plurality of rotatableblades. The rotor is mechanically attached to the shaft. A piston ringis carried by the shaft. The piston ring includes a first layer disposedon a substrate, and a second layer dimensioned to abut against the hubto establish a sealing relationship. The second layer includes a solidlubricant that has molybdenum trioxide (MoO3).

In a further embodiment of any of the foregoing embodiments, the rotoris an integrally bladed rotor.

In a further embodiment of any of the foregoing embodiments, thecompressor section includes a low pressure compressor and a highpressure compressor, and the integrally bladed rotor is a high pressurecompressor rotor of the high pressure compressor.

A method of sealing for a gas turbine engine according to an example ofthe present disclosure includes forming a piston ring including a firstlayer disposed on a substrate, and a second layer disposed on the firstlayer. The second layer includes a solid lubricant that has molybdenumtrioxide (MoO3). The method includes mechanically attaching a rotor to arotatable shaft such that the piston ring establishes a sealingrelationship between the rotor and the shaft. The rotor includes a hubcarrying one or more rotatable blades.

In a further embodiment of any of the foregoing embodiments, the step offorming includes depositing the second layer directly on the firstlayer.

In a further embodiment of any of the foregoing embodiments, the step offorming includes depositing molybdenum disulfide (MoS2) on the firstlayer, and heating the molybdenum disulfide (MoS2) to a predeterminedtemperature threshold to form the solid lubricant.

In a further embodiment of any of the foregoing embodiments, the firstlayer has a composition, by weight percent, 85% to 95% copper, and atleast 5% aluminum.

In a further embodiment of any of the foregoing embodiments, the pistonring includes a seal body extending between a first end and a second,opposed end that engages with the first end along an outer diameterportion of the shaft.

In a further embodiment of any of the foregoing embodiments, the hubdefines a seal land that establishes the sealing relationship with thesecond layer.

In a further embodiment of any of the foregoing embodiments, the rotoris an integrally bladed rotor, and the second layer is dimensioned toabut against an inner diameter portion of the hub to establish thesealing relationship.

The various features and advantages of this disclosure will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a gas turbine engine.

FIG. 2 illustrates a rotor assembly for a gas turbine engine.

FIG. 3 illustrates an enhanced view of a portion of the rotor assemblyof FIG. 2 including a seal.

FIG. 4 illustrates a perspective view of the seal of FIG. 3.

FIG. 5 illustrates a sectional view of the rotor assembly of FIG. 3.

FIG. 6A illustrates an example plot of coefficient of friction (COF)versus cycles of a rotor assembly.

FIG. 6B illustrates an example plot of COF versus cycles of anotherrotor assembly.

FIG. 7 illustrates an example process for forming a rotor assembly.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a housing15 such as a fan case or nacelle, and also drives air along a core flowpath C for compression and communication into the combustor section 26then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 may be arranged generallybetween the high pressure turbine 54 and the low pressure turbine 46.The mid-turbine frame 57 further supports bearing systems 38 in theturbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and less than about 5:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R.)/(518.7° R.)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

FIG. 2 illustrates a rotor assembly 60 for a section 58 of a gas turbineengine, such as the compressor section 24 of FIG. 1. Although thedisclosure primarily refers to the compressor section 24, other portionsof the engine 20 may benefit from the teachings disclosed herein, suchas the fan and turbine sections 22, 28, towershafts and auxiliarysystems. Other systems may also benefit from the teachings herein,including marine systems, and ground-based systems lacking a fan forpropulsion.

The section 58 includes a plurality of rotors 62 each including a diskor hub 63 that carries one or more rotatable blades or airfoils 64. Theairfoils 64 are rotatable about the engine axis A in a gas path GP, suchas core flow path C. Each airfoil 64 includes a platform 64A and anairfoil section 64B extending in a spanwise or radial direction R fromthe platform 64A to a tip 64C. A root section 64D of each airfoil 64extends outwardly from, and is mounted to, a respective hub 63. In someexamples, the root section 64D is received in a slot defined by the hub63. In the illustrated example of FIG. 2, the rotor 62 is a blisk orintegrated bladed rotor (IBR) in which the airfoils 64 are integrallyformed with the hub 63. In examples, the IBR is a compressor rotor, suchas a high pressure compressor rotor of the high pressure compressor 52coupled to the high pressure turbine 54, or such as a low pressurecompressor rotor of the low pressure compressor 44 coupled to the lowpressure turbine 46. Various techniques can be utilized to form the IBR,such as casting, additive manufacturing, machining from a solid workpiece, or welding individual airfoils 64 to the hub 63.

One or more rows of vanes 66 are positioned along the engine axis A andadjacent to the airfoils 64 to direct flow in the gas path GP. The vanes66 can be mechanically attached to the engine static structure 38 (FIG.1). An array of seals 68 are distributed about each row of airfoils 64to bound the gas path GP.

One or more of the rotors 62 are mechanically attached or otherwisefixedly secured to an elongated, rotatable shaft 70. The shaft 70 isrotatable about the engine axis A. In examples, the shaft 70interconnects a turbine and a compressor, such as one of the shafts 40,50 of FIG. 1.

The rotor assembly 60 includes at least one annular seal 74. The seal 74can be located at one or more of the stages of the section 58, such asan intermediate stage as illustrated by seal 74, and/or a forwardmost oraftmost stage indicated by seals 74′, 74″ of rotor assemblies 60′, 60″(74′, 74″ shown in dashed lines for illustrative purposes). Each rotorassembly 60 can have a single seal as illustrated by seals 74, 74′ orcan have multiple seals as illustrated by seals 74″ of rotor assembly60″.

FIG. 3 illustrates an enhanced view of a portion 72 of the rotorassembly 60. The annular seal 74 is carried by, and extends about, anouter diameter portion 70A the shaft 70 for establishing a sealingrelationship between the hub 63 of the rotor 62 and the shaft 70 toblock or otherwise reduce communication of flow between adjacentcavities defined by the rotor 62 and shaft 70. A cross section of theshaft 70 can have a circular or otherwise generally elliptical geometry.

The seal 74 can be a piston ring having a generally hoop-shapedgeometry, as illustrated in FIG. 4. In the illustrated example of FIG.4, the seal 74 is a split ring including a main body 74A extendingbetween opposed first and second ends 74B, 74C. The first end 74Bengages with the second end 74C at an interface along the outer diameterportion 70A of the shaft 70. In other examples, the body 74A iscontinuous to form a full hoop.

The seal 74 is dimensioned to extend about a circumference of the shaft70. The shaft 70 includes an annular groove 76 extending inwardly fromthe outer diameter portion 70A. The groove 76 is dimensioned to receiveat least a portion of the seal 74. The seal 74 is seated in the groove76 to establish a sealing relationship with an inner diameter portion63A of the hub 63. The seal 74 can be arranged such that the first end74B engages with the second end 74C along the outer diameter portion 70Aof the shaft 70. The seal 74 can be exposed to relatively hightemperatures due to proximity to the gas path GP and other portions ofthe engine 20.

Referring to FIG. 5, with continuing reference to FIGS. 3-4, the seal 74has a multi-layer construction including a substrate 78 and a coating 80including a plurality of layers. The coating 80 can include at least afirst layer 80A and a second layer 80B. At least a portion of thesubstrate 78 is seated in the groove 76, and the coating 80 is situatedbetween the substrate 78 and the hub 63. The second layer 80B can definean external surface of the seal 74. The inner diameter portion 63A ofthe hub 63 defines a counter face or seal land. The second layer 80B canbe dimensioned to abut against the hub 63 to establish a sealingrelationship with the seal land. In the illustrative example of FIG. 5,the second layer 80B is dimensioned and arranged to abut against orcontact the inner diameter portion 63A of the hub 63, or is otherwise inclose proximity to the seal land, to establish the sealing relationshipwith the hub 63 along an interface 84.

The first layer 80A can be disposed directly on the substrate 78. Thesecond layer 80B can be disposed directly on the first layer 80A. Inother examples, one or more layers of material are formed between thesubstrate 78 and first layer 80A and/or between the first and secondlayers 80A, 80B. The hub 63 can include a third layer 86 (shown indashed lines for illustrative purposes) disposed on surfaces of theinner diameter portion 63A along the interface 84 to establish the sealland. In other examples, the third layer 86 is omitted.

Various materials can be utilized to form the rotor assembly 60. Thefirst layer 80A can be made of a first material, the second layer 80Bcan be made of a second material, and the substrate 78 can be made of athird material. The first, second and/or third materials can differ incomposition and/or construction.

The substrate 78 and/or rotor 62 can comprise a high temperature metalor metal alloy, such as a nickel alloy. Example nickel alloys includenickel chromium alloy sold under the tradename INCONEL® alloy 718(IN718) and Direct Age Processed Alloy 718 (DA718). In the illustratedexample of FIG. 5, the substrate 78 is made of IN718 and at least theinner diameter portion 63A of the hub 63 is made of DA718. In otherexamples, the substrate 78 comprises a cobalt alloy sold under thetradename STELLITE® 6B, and the first layer 80A can be omitted.

The first layer 80A may be a relatively soft metal or metal alloycoating comprising copper or a copper alloy including aluminum ornickel, for example. In other examples, the first layer 80A is anickel-based or molybdenum-based metal or metal alloy. In examples, thefirst layer 80A has a composition, by weight percent, of about 60% toabout 95% copper, and aluminum. In further examples, the first layer 80Ahas a composition, by weight percent, of about 85% to about 95% copper,and at least 5% aluminum. In examples, the first layer 80A has acomposition, by weight percent, 8% to 12% aluminum, and at least 85%copper. In further examples, the first layer 80A has a composition, byweight percent, of about 90% copper, and about 10% aluminum. In otherexamples, the first layer 80A has a composition of copper including anyof the weight percentages disclosed herein, and the balance is aluminumor nickel. For the purposes of this disclosure, the term “about” means±3% of the disclosed weight percent value unless otherwise stated.

The second layer 80B can be a low friction coating comprising a solidlubricant. The second layer 80B can have a lesser hardness than thefirst layer 80A. Hardness of the layers 80A, 80B can be measured bymeans of nanoindentation.

In the illustrated example of FIG. 5, the first layer 80A of the coating80 is a layer of copper alloy disposed on the substrate 78, and thesecond layer 80B is a solid lubricant disposed on the first layer 80A. Ahardness of the copper-based coating can be relatively lower than othermaterials to reduce wear of the counter face and the solid lubricantthat may otherwise occur due to the interface 84. The second layer 80Bprovides a self-lubricating feature and can reduce frictional heatingalong the interface 84.

The solid lubricant of the second layer 80B can comprise molybdenumtrioxide (MoO3), for example. The second layer 80B can be formedutilizing any of the techniques disclosed herein, including applying asolid lubricant such as molybdenum disulfide (MoS2) on the first layer80A or substrate 78 and heat treating the solid lubricant to form MoO3.The solid lubricant can reduce a coefficient of friction (COF) betweenthe seal 74 and the rotor 62 along the interface 84, which can reducewear and improve durability of the components of the rotor assembly 60,including the seal 74 and rotor 62. The second layer 80B can have arelatively lower COF than the first layer 80A of the coating 80.

The third layer 86 of the hub 63 can be a solid lubricant, which can bethe same or can differ from the solid lubricant of the second layer 80B.The solid lubricant of the third layer 86 can comprise MoS2, MoO3, boronnitride, tungsten disulfide, and carbon-based and graphite-basedmaterials, for example. Applying a solid lubricant to the inner diameterportion 63A of the hub 63 can reduce leakage along the interface 84.

During operation, the shaft 70 and each rotor 62 rotate as an assemblyabout the engine axis A. The seal 74 may move relative to the hub 63along the interface 84 in axial, radial and/or circumferentialdirections X, R, C. The axial direction X can be coincidental orparallel to the engine axis A. Sliding or movement of the seal 74 alongthe interface 84 in the axial and/or circumferential directions X, C mayoccur due to relatively high vibratory energy in the system. The solidlubricant can reduce the COF and frictional heating along the interface84, which can reduce galling and other wear along the adjacent surfacesand can increase durability of the seal 74 and the rotor 62. Reducedwear can reduce overhaul costs that may otherwise be associated withreplacement or refurbishment of the seal 74, shaft 70 and/or rotor 62.

FIGS. 6A and 6B illustrate example plots of COF versus cycles of rotorassemblies. The x-axis corresponds to the number of cycles of movementof the respective seal relative to the interface. The y-axis correspondsto the average COF. Values along the y-axis correspond to movement ofthe seal in a first direction. FIG. 6A illustrates curve 88corresponding to average COF values for a seal having a copper-basedcoating free of a solid lubricant. FIG. 6B illustrates curve 90corresponding to average COF values for the seal 74 having acopper-based coating and a MoO3-based solid lubricant. The MoO3-basedsolid lubricant can be formed from a MoO3-based solid lubricantutilizing any of the techniques disclosed herein.

As illustrated by FIGS. 6A-6B, the copper-based coating including aMoO3-based solid lubricant can reduce the average COF as compared tocoatings free of a solid lubricant. In the illustrative example of FIG.6B, the seal 74 has an average COF of less than 0.4, such as less than0.3 when the solid lubricant is first formed and between 0.3 and 0.35subsequent to an initial break-in period of the seal 74, whereas theaverage COF of curve 88 is greater than 0.5 for a comparable number ofcycles. The combination of materials of the substrate 78, layers 80A,80B of coating 80 and/or hub 63 can reduce wear adjacent the first andsecond ends 74B, 74C of the seal 74 and adjacent portions of the hub 63.

FIG. 7 illustrates a process in a flowchart 92 for forming a rotorassembly, including any of the seals and rotor assemblies disclosedherein. Reference is made to the seal 74 and rotor assembly 60 of FIGS.3-5 for illustrative purposes. In examples, a coating including one ormore layers of material is removed from surfaces of substrate 78 at step94, including at least a previously applied second layer 80B comprisedof a solid lubricant.

The first layer 80A is deposited or is otherwise disposed or formed onthe substrate 78 at step 95. The first layer 80A can be formed on thesubstrate 78 utilizing various techniques, such as by plasma spraydeposition or another thermal spraying technique. Example techniques forforming the first layer 80A can include physical vapor deposition (PVD)and chemical vapor deposition (CVD). In examples, step 94 includesremoving a previously applied first layer 80A such as a layer of copperalloy from the substrate 78 prior to step 95. The previously appliedfirst layer 80A may be removed after about 10,000-12,000 operatingcycles of the engine, whereas a previously applied second layer 80B maybe removed after a lesser number of operating cycles, such asapproximately 1000 operating cycles, for example.

At step 96 the second layer 80B comprising a solid lubricant such asMoO3 is disposed or otherwise formed on the first layer 80A. Step 96 caninclude depositing the second layer 80B directly on the first layer 80A.In some examples, at step 99 the third layer 86 comprising a solidlubricant is deposited or otherwise formed on the hub 63. A previouslyapplied third layer 86 can be removed prior to steps 95, 96.

The second and third layers 80B, 86 can be applied utilizing varioustechniques, such as brushing, swabbing, or spraying including any of thespraying techniques disclosed herein. In examples, the second and/orthird layers 80B, 86 are applied by PVD or CVD techniques.

Step 96 can include depositing a precursor on the first layer 80A atstep 97, and heating the precursor to a predetermined temperaturethreshold for a predetermined time threshold to form the solid lubricantat step 98. In examples, the precursor is a solid lubricant such as MoS2and is applied by brushing, swabbing or spraying to obtain a thicknessof about 0.0001-0.001 inches at step 97, for example, and step 98includes forming a solid lubricant comprising MoO3 in response toheating the precursor to a predetermined temperature threshold of atleast about 750° F., including oxidation of the precursor such as MoS2.The MoS2 or other solid lubricant can include a binder such as silicate.The MoS2 applied to the first layer 80A at step 97 can be free of anyMoO3. Forming the MoO3 from MoS2 can reduce manufacturing cost andcomplexity. In examples, step 98 includes pre-treating the coating 80 byheating the MoS2 deposited on the substrate 78 or first layer 80A toabout 930° F. for about 8 hours to form the second layer 80B of MoO3.The solid lubricant may comprise no more than about 10% of binderconstituents such that a majority, such as at least 90%, of the secondlayer 80B is MoO3. The predetermined time threshold can be greater thanabout 8 hours to increase a concentration of the MoO3 forming the secondlayer 80B.

The first, second and/or third layers 80A, 80B, 86 may include one ormore sublayers. One or more layers of material may be formed on and/orbetween the layers 80A, 80B, 86. Various finishing operations can beapplied to the seal 74 once the respective layers 80A, 80B, 86 areformed.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A rotor assembly for a gas turbine enginecomprising: a rotor including a hub carrying one or more rotatableblades, the rotor mechanically attached to a shaft; and an annular sealcarried by the shaft, wherein the annular seal comprises: a substrate; afirst layer disposed on the substrate; and a second layer disposed onthe first layer and arranged to establish a sealing relationship withthe rotor, the second layer comprising a solid lubricant includingmolybdenum trioxide (MoO3).
 2. The rotor assembly as recited in claim 1,wherein the first layer comprises copper.
 3. The rotor assembly asrecited in claim 2, wherein the first layer has a composition, by weightpercent, 60% to 95% copper, and aluminum.
 4. The rotor assembly asrecited in claim 2, wherein the first layer has a composition, by weightpercent, 8% to 12% aluminum, and at least 85% copper.
 5. The rotorassembly as recited in claim 1, wherein the substrate comprises a nickelalloy.
 6. The rotor assembly as recited in claim 1, wherein the rotor isan integrally bladed rotor, and the annular seal extends about an outerdiameter portion of the shaft.
 7. The rotor assembly as recited in claim6, wherein the first layer has a composition, by weight percent, 85% to95% copper, and at least 5% aluminum.
 8. The rotor assembly as recitedin claim 6, wherein the annular seal is a split ring including a sealbody extending between a first end and a second, opposed end thatengages with the first end.
 9. The rotor assembly as recited in claim 6,wherein the integrally bladed rotor is a compressor rotor, and thesubstrate of the annular seal is seated in an annular groove extendinginwardly from the outer diameter portion of the shaft, and the secondlayer abuts against an inner diameter portion of the hub.
 10. The rotorassembly as recited in claim 1, wherein the molybdenum trioxide (MoO3)is formed from molybdenum disulfide (MoS2).
 11. A gas turbine enginecomprising: a fan section including a fan; a compressor sectionincluding a compressor; a turbine section including a turbine coupled tothe compressor; a shaft rotatable about an engine longitudinal axis; andwherein the compressor section includes a rotor assembly, the rotorassembly comprising: a rotor including a hub carrying a plurality ofrotatable blades, the rotor mechanically attached to the shaft; a pistonring carried by the shaft, the piston ring comprising a first layerdisposed on a substrate, and a second layer dimensioned to abut againstthe hub to establish a sealing relationship, and the second layercomprising a solid lubricant including molybdenum trioxide (MoO3). 12.The gas turbine engine as recited in claim 11, wherein the rotor is anintegrally bladed rotor.
 13. The gas turbine engine as recited in claim12, wherein the compressor section includes a low pressure compressorand a high pressure compressor, and the integrally bladed rotor is ahigh pressure compressor rotor of the high pressure compressor.
 14. Amethod of sealing for a gas turbine engine comprising: forming a pistonring comprising a first layer disposed on a substrate, and a secondlayer disposed on the first layer, the second layer comprising a solidlubricant including molybdenum trioxide (MoO3); and mechanicallyattaching a rotor to a rotatable shaft such that the piston ringestablishes a sealing relationship between the rotor and the shaft, therotor including a hub carrying one or more rotatable blades.
 15. Themethod as recited in claim 14, wherein the step of forming includesdepositing the second layer directly on the first layer.
 16. The methodas recited in claim 14, wherein the step of forming includes depositingmolybdenum disulfide (MoS2) on the first layer, and heating themolybdenum disulfide (MoS2) to a predetermined temperature threshold toform the solid lubricant.
 17. The method as recited in claim 16, whereinthe first layer has a composition, by weight percent, 85% to 95% copper,and at least 5% aluminum.
 18. The method as recited in claim 16, whereinthe piston ring includes a seal body extending between a first end and asecond, opposed end that engages with the first end along an outerdiameter portion of the shaft.
 19. The method as recited in claim 14,wherein the hub defines a seal land that establishes the sealingrelationship with the second layer.
 20. The method as recited in claim14, wherein: the rotor is an integrally bladed rotor; and the secondlayer is dimensioned to abut against an inner diameter portion of thehub to establish the sealing relationship.